This invention relates to nickel-base superalloys and, more particularly, to such a superalloy optimized for use in high-temperature components of jet engines such as turbine disks.
In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on the same shaft. The turbine includes a disk portion mounted to the shaft and a series of turbine blades supported on the rim of the disk. The flow of hot exhaust gas impinges upon the turbine blades and causes the turbine disk to turn, which turns the shaft and provides power to the compressor. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forwardly.
The hotter the exhaust gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the exhaust combustion gas temperature, which in turn leads to higher operating temperature requirements of many of the components from which the engine is constructed. In response to these requirements, alloys with carefully tailored, improved mechanical properties have been developed for use in the various sections of the engines.
In operation, the turbine disks encounter different operating conditions radially from the center or hub portion to the exterior or rim portion. The rim is hotter than the hub and, in general, all of the operating temperatures are higher for more advanced engines. The stress conditions also vary radially, with the lower stresses at the rim and the higher stresses at the hub. As a result of the different operating conditions, the material at the rim of the disk must exhibit good high temperature creep and stress rupture resistance as well as high-temperature strength and hold-time fatigue crack growth resistance. The hub region of the disk must exhibit high tensile strength at more moderate temperatures and resistance to low cycle fatigue crack growth. In the most common design, the entire turbine disk is made of a single forged and heat-treated piece of material. The selected alloy used in the disk must therefore meet all of the materials requirements discussed above.
The materials used in the turbine disk are also chosen in relation to the aircraft mission requirements. In general, the mission cycles of high-performance military aircraft engines require higher operating temperatures but have shorter. times at the maximum temperatures, as compared with those of civilian aircraft engines. A current goal in some military aircraft applications is a high-pressure turbine disk operable at temperatures of up to 1500xc2x0 F. for relatively short periods of time.
Current nickel-base superalloys used in turbine disks, such as Rene 88DT, Rene 95, and IN 100, have an operating temperature limit of 1200-1300xc2x0 F. These types of alloys cannot meet the operating temperature goal of 1500xc2x0 F. for the military aircraft engines. New alloys are under development which have operating temperature limits approaching about 1400xc2x0 F. under some mission cycles, but typically such alloys have high gamma-prime solvus temperatures and are accordingly difficult to process. Some have been observed to exhibit undesirable thermally induced porosity.
Thus, although satisfactory alloys are available for use in turbine disks for existing engines and there are development efforts underway for alloys with evenhigher operating temperatures, there is always a need for improved materials that are operable in applications such as aircraft turbine disks at higher temperatures of up to 1500xc2x0 F., are stable, and are producible. The present invention provides such an improved material.
The present invention provides a nickel-base superalloy composition that is useful in hot-section components of aircraft gas turbine engines. The alloy is particularly useful in turbine disks for the high-pressure turbine stages of the engine that are subjected to the highest operating temperatures. The alloy is optimized for superior mechanical performance in operating cycles reaching 1500xc2x0 F., and is also selected for good fabrication and producibility properties. The density of the alloy is about 0.301 pounds per cubic inch, which is acceptable and does not lead to overly high centrifugal stresses during service. Alloy phase stability and chemical stability are good, an important consideration for an alloy which is to be used at temperatures as high as 1500xc2x0 F., even for relatively short times.
In accordance with the invention, a composition of matter comprises in combination, in weight percent, from about 16.0 percent to about 22.4 percent cobalt, from about 6.6 percent to about 14.3 percent chromium, from about 1.4 percent to about 3.5 percent tantalum, from about 1.9 percent to about 4.0 percent tungsten, from about 1.9 percent to about 3.9 percent molybdenum, from about 0.03 percent to about 0.10 percent zirconium, from about 0.9 percent to about 3.0 percent niobium, from about 2.4 percent to about 4.6 percent titanium, from about 2.6 percent to about 4.8 percent aluminum, from 0 to about 2.5 percent rhenium, from about 0.02 percent to about 0.10 percent carbon, from about 0.02 percent to about 0.10 percent boron, balance nickel and minor amounts of impurities. Optionally, the following elements may also be present: from 0 to about 2 percent vanadium, from 0 to about 2 percent iron, from 0 to about 2 percent hafnium, and from 0 to about 0.1 percent magnesium.
A preferred composition comprises from about 16.0 percent to about 20.2 percent cobalt, from about 6.6 percent to about 12.5 percent chromium, from about 1.5 percent to about 3.5 percent tantalum, from about 2.0 percent to about 4.0 percent tungsten, from about 1.9 percent to about 3.9 percent molybdenum, from about 0.04 percent to about 0.06 percent zirconium, from about 1.0 percent to about 3.0 percent niobium, from about 2.6 percent to about 4.6 percent titanium, from about 2.6 percent to about 4.6 percent aluminum, from 0 to about 2.5 percent rhenium, from about 0.02 percent to about 0.04 percent carbon, from about 0.02 percent to about 0.04 percent boron, balance nickel and minor amounts of impurities.
These alloys and their most preferred embodiments are carefully optimized for excellent creep performance in turbine disks operating at temperatures approaching 1500xc2x0 F. Dwell fatigue crack growth performance is good, and in some cases excellent, but the primary emphasis is on obtaining the good creep performance required in this operating temperature range. The dwell fatigue crack growth performance is relatively less important than creep performance because of the relatively shorter operating times spent at the maximum elevated temperature in high-performance military engines as compared with civilian engines, for example. Some low-temperature dwell fatigue crack growth performance is therefore intentionally sacrificed in the optimized alloy of the invention to achieve further improved creep performance. The present alloys also achieve a reduced gamma-prime solvus temperature that provides a wider temperature range for heat treatments between the gamma-prime solvus and the solidus temperatures. This wider temperature range improves the processibility of the alloy. The grain boundary elements aid in the retention of a desired grain size.
Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings, which illustrate, by way of example, the principles of the invention. The scope of the invention is not, however, limited to this preferred embodiment.